Turbine rotor blade with integral impingement sleeve by additive manufacture

ABSTRACT

A turbine rotor blade is additively manufactured and includes an airfoil body with a radially extending chamber for receiving a coolant flow, a tip end at a radial outer end of the airfoil body, and a shank at a radial inner end of the airfoil body. The radially extending chamber extends at least partially into the shank to define a shank inner surface. An integral impingement cooling structure is within the radially extending chamber. The integral impingement cooling structure allows an exterior surface of a hollow body thereof to be uniformly spaced from the airfoil inner surface despite the curvature of the chamber. The turbine rotor blade has impingement cooling throughout the blade.

BACKGROUND

The disclosure relates generally to turbomachines, and moreparticularly, to a turbine rotor blade with a number of additivelymanufactured, integral features such as an integral impingement sleeve,non-linear cooling passages in a platform, angel wings with a coolanttransfer passage, and a hollow dovetail with an integral lattice supportstructure.

Turbomachines include a plurality of turbine rotor blades coupled to arotor. A working fluid such as steam or a combusted fuel is forcedagainst the blades to force them to turn the rotor. Turbine rotor bladesoperate in extremely hot conditions and require cooling. Coolingfeatures can be provided in a number of ways.

One mechanism to provide cooling is an impingement insert. Animpingement insert or sleeve includes a hollow body having coolingpassages in a wall thereof that allow delivery of a coolant through thecooling passages to impact or impinge on a surface to be cooled.Impingement inserts are used, for example, in a variety of hot gas path(HGP) components in turbomachinery such as a turbine rotor blade toincrease cooling performance of cooling circuitry therein. One challengewith impingement inserts is positioning the impingement insert within atapered or curved cavity in an HGP component in a sufficiently closemanner to allow for high cooling performance, but not so close thatcooling is ineffective. One indicator of cooling performance of animpingement insert is the Z/D parameter, which is a ratio of a standoffdistance Z between the insert and interior surface of the HGP componentand a diameter D of the cooling passages (holes) in the impingementinsert. The Z/D parameter value of an insert is typically designed to bewithin a desired range that results in better cooling performance.

Impingement cooling is typically not provided by an insert where thenecessary standoff distance cannot be created. For example, where thecavity in the HGP component curves too significantly that an impingementinsert cannot be made sufficiently thin or curved to respect thenecessary standoff distance, impingement cooling cannot be provided. Oneapproach to address this challenge provides the impingement insert in anumber of flexible, longitudinal sections to make it easier to insertthem into the HGP component. However, having to sequentially positionand couple a number of insert sections together or couple them to theHGP component, increases the complexity, time and costs of manufacture.Flexible impingement insert sections also do not provide a contiguouselement about their periphery, i.e., laterally (cross-section), whichcan detract from cooling performance where they are discontinuous.

Impingement cooling has been applied in a limited manner to rotatingturbine rotor blades in turbomachines, e.g., for leading edges thereof.However, impingement cooling has not been applied more broadly across anentirety of an inner surface of a turbine rotor blades because thecentrifugal forces experienced by the rotating blades forces the coolantto the radially outer tip end of the blade as it rotates, makingimpingement cooling less effective.

Another cooling feature includes cooling passages through a part of theturbine rotor blade to be cooled. For example, the turbine rotor bladesinclude a platform that extends laterally to form part of a workingfluid path through the turbomachine in cooperation with a platform of anadjacent turbine rotor blade. Due to the high temperatures of theworking fluid, the platforms typically include a cooling circuit thereinthat feeds a number of cooling passages that exit through a slash faceof the platform. Some platforms include a damping pin seat in the slashface that receives an axially extending pin therein that mates with anadjacent damping pin seat in an adjacent platform to seal the workingfluid path. The cooling passages are typically drilled into the slashface to fluidly couple the passages to the cooling circuit.Consequently, the cooling passages have a linear configuration that maynot adequately cool all of the platform. For example, the coolingpassages may pass through an extension that forms a damping pin seat,but inadequately cool other portions of the slash face.

Cooling features may also be employed with angel wings. In this regard,another cooling feature includes cooling passages that either delivercoolant into the angel wing, or radially around an angel wing. Mountsfor turbine rotor blades may also include cooling features therein.

BRIEF DESCRIPTION

A first aspect of the disclosure provides a turbine rotor blade,comprising: an airfoil body including a concave pressure side outer walland a convex suction side outer wall that connect along leading andtrailing edges, the outer walls having an airfoil inner surface defininga radially extending chamber for receiving a coolant flow; a tip end ata radial outer end of the airfoil body; a shank at a radial inner end ofthe airfoil body, the radially extending chamber extending at leastpartially into the shank to define a shank inner surface; and animpingement cooling structure within the radially extending chamber, theimpingement cooling structure including: a hollow body including a firstend, a second end, an interior surface and an exterior surface, aplurality of cooling passages through the hollow body and in fluidcommunication with the radially extending chamber to allow the coolantflow to pass from the interior surface of the hollow body to impinge onat least the airfoil inner surface, wherein the first end of the hollowbody is integrally formed to the shank inner surface, and wherein theexterior surface of the hollow body is uniformly spaced from the airfoilinner surface between the first end and the second end of the hollowbody.

A second aspect of the disclosure provides an additively manufacturedturbine rotor blade, comprising: an airfoil body including a concavepressure side outer wall and a convex suction side outer wall thatconnect along leading and trailing edges, the outer walls having anairfoil inner surface defining a radially extending chamber forreceiving a coolant flow; and an integral impingement cooling structurewithin the radially extending chamber, the integral impingement coolingstructure including: a hollow body including a first end, a second end,an interior surface and an exterior surface, and a plurality of coolingpassages through the hollow body and in fluid communication with theradially extending chamber to allow the coolant flow to pass from theinterior surface of the hollow body to impinge on at least the airfoilinner surface, wherein the exterior surface of the hollow body isuniformly spaced from the airfoil inner surface between the first endand the second end of the hollow body.

A third aspect of the disclosure provides a method comprising:sequentially creating a layer of material and applying a heat source tosinter the layer of materials to form: an airfoil body including aconcave pressure side outer wall and a convex suction side outer wallthat connect along leading and trailing edges, the outer walls having anairfoil inner surface defining a radially extending chamber forreceiving a coolant flow; and an impingement cooling structure withinthe radially extending chamber, the integral impingement coolingstructure including: a hollow body including a first end, a second end,an interior surface and an exterior surface, and a plurality of coolingpassages through the hollow body and in fluid communication with theradially extending chamber to allow the coolant flow to pass from theinterior surface of the hollow body to impinge on at least the airfoilinner surface, wherein the exterior surface of the hollow body isuniformly spaced from the airfoil inner surface between the first endand the second end of the hollow body.

A fourth aspect of the disclosure provides a turbine rotor blade,comprising: an airfoil body including a concave pressure side outer walland a convex suction side outer wall that connect along leading andtrailing edges, the outer walls defining a radially extending chamberfor receiving a coolant flow; a platform extending laterally outwardrelative to the airfoil body and terminating at at least one slash face;a cooling circuit defined within the platform and in fluid communicationwith a source of the coolant flow; and at least one cooling passagedefined in the platform and in fluid communication with the coolingcircuit, the at least one cooling passage extending in a non-linearconfiguration from the cooling circuit to exit through the at least oneslash face of the platform.

A fifth aspect of the disclosure provides an additively manufacturedturbine rotor blade, comprising: an airfoil body including a concavepressure side outer wall and a convex suction side outer wall thatconnect along leading and trailing edges, the outer walls defining aradially extending chamber for receiving a coolant flow; a platformextending laterally outward relative to the airfoil body and terminatingat least one slash face; a cooling circuit defined within the platformand in fluid communication with a source of the coolant flow; and atleast one cooling passage defined in the platform and in fluidcommunication with the cooling circuit, the at least one cooling passageextending in a non-linear configuration from the cooling circuit to exitthrough the slash face of the platform.

A sixth aspect includes a turbine rotor blade, comprising: an airfoilbody including a concave pressure side outer wall and a convex suctionside outer wall that connect along leading and trailing edges; a shankat a radial inner end of the airfoil body; at least one angel wingextending laterally from at least one side of the shank; and a coolanttransfer passage defined through the at least one angel wing, thecoolant transfer passage fluidly coupling a first wheel space portiondefined between the shank and a first adjacent shank of a first adjacentturbine rotor blade and a second wheel space portion defined between theshank and a second adjacent shank of a second adjacent turbine rotorblade.

A seventh aspect of the disclosure relates to an additively manufacturedturbine rotor blade, comprising: an airfoil body including a concavepressure side outer wall and a convex suction side outer wall thatconnect along leading and trailing edges; a shank at a radial inner endof the airfoil body; at least one angel wing extending laterally from atleast one side of the shank; and a coolant transfer passage definedthrough the at least one angel wing, the coolant transfer passagefluidly coupling a first wheel space portion defined between the shankand a first adjacent shank of a first adjacent turbine rotor blade and asecond wheel space portion defined between the shank and a secondadjacent shank of a second adjacent turbine rotor blade.

An eighth aspect relates to a set of turbine rotor blades, comprising: afirst turbine rotor blade, a second turbine rotor blade and a thirdturbine rotor blade, the first turbine rotor blade positioned betweenthe second and third turbine rotor blades, each turbine rotor bladeincluding: an airfoil body including a concave pressure side outer walland a convex suction side outer wall that connect along leading andtrailing edges, a shank at a radial inner end of the airfoil body, atleast one angel wing extending laterally from at least one side of theshank, and wherein the shanks of the first and second turbine rotorblades define a first wheel space portion therebetween and the shanks ofthe first and third turbine rotor blades define a second wheel spaceportion therebetween; and a coolant transfer passage defined through theat least one angel wing in the first turbine rotor blade, the coolanttransfer passage fluidly coupling the first wheel space portion and thesecond wheel space portion.

A ninth aspect relates to a turbine rotor blade root, comprising: ashank having a radially extending chamber defined therein; a blade mountat a radial inner end of the shank, the blade mount having a hollowinterior defined therein, the hollow interior in fluid communicationwith the radially extending chamber; and a lattice support structuredisposed within the hollow interior of the blade mount.

A tenth aspect includes a turbine rotor blade root, comprising: a shankhaving a radially extending chamber defined therein; a blade mount at aradial inner end of the shank, the blade mount having a hollow interiordefined therein, the hollow interior in fluid communication with theradially extending chamber; a lattice support structure disposed withinthe hollow interior of the blade mount; at least one angel wingextending laterally from at least one side of the shank; and a coolanttransfer passage defined through the at least one angel wing, thecoolant transfer passage fluidly coupling a first wheel space portiondefined between the shank and a first adjacent shank of a first adjacentturbine rotor blade root and a second wheel space portion definedbetween the shank and a second adjacent shank of a second adjacentturbine rotor blade root.

An eleventh aspect relates to a turbine rotor blade root, comprising: ashank having a radially extending chamber defined therein; a blade mountat a radial inner end of the shank, the blade mount having a hollowinterior defined therein, the hollow interior in fluid communicationwith the radially extending chamber; a lattice support structuredisposed within the hollow interior of the blade mount; at least oneangel wing extending laterally from at least one side of the shank; anda coolant transfer passage defined through the at least one angel wing,the coolant transfer passage fluidly coupling a first wheel spaceportion defined between the shank and a first adjacent shank of a firstadjacent turbine rotor blade root and a second wheel space portiondefined between the shank and a second adjacent shank of a secondadjacent turbine rotor blade root.

The illustrative aspects of the present disclosure are designed to solvethe problems herein described and/or other problems not discussed.

BRIEF DESCRIPTION OF THE DRAWINGS

These and other features of this disclosure will be more readilyunderstood from the following detailed description of the variousaspects of the disclosure taken in conjunction with the accompanyingdrawings that depict various embodiments of the disclosure, in which:

FIG. 1 shows a schematic view of an illustrative gas turbine (GT)system.

FIG. 2 shows a cross-sectional view of an illustrative gas turbineassembly that may be used with the GT system in FIG. 1.

FIG. 3 shows a perspective view of a turbine rotor blade of the type inwhich embodiments of the present disclosure may be employed.

FIG. 4 shows an axial cross-sectional view of a turbine rotor bladeincluding an integral impingement sleeve, according to one embodiment ofthe disclosure.

FIG. 5 shows a radial, circumferential cross-sectional view of a turbinerotor blade including an integral impingement sleeve, according to oneembodiment of the disclosure.

FIG. 6 shows an enlarged cross-sectional view of a meeting location ofan impingement cooling structure and a shank of the turbine rotor blade,according to another embodiment of the disclosure.

FIG. 7 shows an enlarged cross-sectional view of an impingement coolingstructure with a variable wall thickness, according to anotherembodiment of the disclosure.

FIG. 8 shows an enlarged cross-sectional view of an impingement coolingstructure with a reinforcement member about a cooling passage thereof,according to an embodiment of the disclosure.

FIG. 9 shows a plan, cross-sectional view of an impingement coolingstructure, according to alternative embodiments of the disclosure.

FIG. 10 shows a first partial axial cross-sectional view of a turbinerotor blade including an integral impingement sleeve, according to oneembodiment of the disclosure.

FIG. 11 shows a second partial axial cross-sectional view of a turbinerotor blade including an integral impingement sleeve, according to oneembodiment of the disclosure.

FIG. 12 shows a perspective view of a cooling passage in a platform of aturbine rotor blade, according to embodiments of the disclosure.

FIG. 13 shows a transparent plan view of a cooling passage in a side ofa platform of a turbine rotor blade, according to embodiments of thedisclosure.

FIG. 14 shows an enlarged cross-sectional view of a cooling passage in aplatform of a turbine rotor blade, according to embodiments of thedisclosure.

FIG. 15 shows an enlarged cross-sectional view of a cooling passage in aplatform of a turbine rotor blade, according to embodiments of thedisclosure.

FIG. 16 shows an enlarged cross-sectional view of a cooling passage in aplatform of a turbine rotor blade, according to embodiments of thedisclosure.

FIG. 17 shows an enlarged cross-sectional view of a cooling passage in aplatform of a turbine rotor blade, according to embodiments of thedisclosure.

FIG. 18 shows an enlarged cross-sectional view of a cooling passage in aplatform of a turbine rotor blade, according to embodiments of thedisclosure.

FIG. 19 shows a cross-sectional view of a turbine rotor blade includingan angel wing, according to embodiments of the disclosure.

FIG. 20 shows a transparent, perspective view of a turbine rotor bladeincluding an angel wing, according to embodiments of the disclosure.

FIG. 21 shows a schematic axial view of a set of turbine rotor bladesincluding an angel wing, according to embodiments of the disclosure.

FIG. 22 shows a plan view a turbine rotor blade including an angel wing,according to embodiments of the disclosure.

FIG. 23 shows a side view of a turbine rotor blade including an angelwing, according to embodiments of the disclosure.

FIG. 24 shows a side view of a turbine rotor blade including an angelwing, according to embodiments of the disclosure.

FIG. 25 shows a cross-sectional view of a hollow blade mount of aturbine rotor blade including a lattice support structure, according toembodiments of the disclosure.

FIG. 26 shows a perspective, cross-sectional view of a root of a turbinerotor blade including a lattice support structure, according toembodiments of the disclosure

It is noted that the drawings of the disclosure are not necessarily toscale. The drawings are intended to depict only typical aspects of thedisclosure, and therefore should not be considered as limiting the scopeof the disclosure. In the drawings, like numbering represents likeelements between the drawings.

DETAILED DESCRIPTION

As an initial matter, in order to clearly describe the currentdisclosure it will become necessary to select certain terminology whenreferring to and describing relevant machine components within, forexample, a turbomachine. When doing this, if possible, common industryterminology will be used and employed in a manner consistent with itsaccepted meaning. Unless otherwise stated, such terminology should begiven a broad interpretation consistent with the context of the presentapplication and the scope of the appended claims. Those of ordinaryskill in the art will appreciate that often a particular component maybe referred to using several different or overlapping terms. What may bedescribed herein as being a single part may include and be referenced inanother context as consisting of multiple components. Alternatively,what may be described herein as including multiple components may bereferred to elsewhere as a single part.

In addition, several descriptive terms may be used regularly herein, andit should prove helpful to define these terms at the onset of thissection. These terms and their definitions, unless stated otherwise, areas follows. As used herein, “downstream” and “upstream” are terms thatindicate a direction relative to the flow of a fluid, such as theworking fluid through the turbomachine or, for example, the flow of airthrough the combustor or coolant through one of the turbine'scomponents. The term “downstream” corresponds to the direction of flowof the fluid, and the term “upstream” refers to the direction oppositeto the flow. The terms “forward” and “aft,” without any furtherspecificity, refer to directions, with “forward” referring to the frontend of the turbomachine (i.e., compressor end) or a component thereof,and “aft” referring to the rearward end of the turbomachine (i.e.,turbine end) or component thereof. Forward and aft generally denoted byan X direction in the drawings. It is often required to describe partsthat are at differing radial positions with regard to a center axis. Theterm “radial” refers to movement or position perpendicular to an axis,e.g., a turbomachine rotor axis. In cases such as this, if a firstcomponent resides closer to the axis than a second component, it will bestated herein that the first component is “radially inward” or “inboard”of the second component. If, on the other hand, the first componentresides further from the axis than the second component, it may bestated herein that the first component is “radially outward” or“outboard” of the second component. Radial direction is generallydenoted by a Z direction in the drawings. The term “axial” refers tomovement or position parallel to an axis, i.e., a turbomachine rotoraxis. Finally, the term “circumferential” refers to movement or positionaround an axis. Although not shown as curved in the legends in thedrawings, the circumferential direction is generally denoted by a Ydirection in the drawings. It will be appreciated that such terms may beapplied in relation to the rotor axis of a turbomachine.

In addition, several descriptive terms may be used regularly herein, asdescribed below. The terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terminology used herein is for the purpose of describing particularembodiments only and is not intended to be limiting of the disclosure.As used herein, the singular forms “a”, “an” and “the” are intended toinclude the plural forms as well, unless the context clearly indicatesotherwise. It will be further understood that the terms “comprises”and/or “comprising,” when used in this specification, specify thepresence of stated features, integers, steps, operations, elements,and/or components, but do not preclude the presence or addition of oneor more other features, integers, steps, operations, elements,components, and/or groups thereof. “Optional” or “optionally” means thatthe subsequently described event or circumstance may or may not occur,and that the description includes instances where the event occurs andinstances where it does not.

Where an element or layer is referred to as being “on,” “engaged to,”“connected to,” or “coupled to” another element or layer, it may bedirectly on, engaged, connected or coupled to the other element orlayer, or intervening elements or layers may be present. In contrast,when an element is referred to as being “integral,” “directly on,”“directly engaged to,” “directly connected to,” or “directly coupled to”another element or layer, there may be no intervening elements or layerspresent. Other words used to describe the relationship between elementsshould be interpreted in a like fashion (e.g., “between” versus“directly between,” “adjacent” versus “directly adjacent,” etc.). Asused herein, the term “and/or” includes any and all combinations of oneor more of the associated listed items.

As indicated above, the disclosure provides a turbine rotor blade orturbine rotor blade root including a number of integral features whichare made possible through additive manufacturing of the blade and/orroot. The additive manufacturing allows formation of structures thatprovide cooling where not previously allowed, improve cooling comparedto conventional systems, provide additional structural strength and/orlower the weight of the blade.

A. Introduction

FIG. 1 shows a schematic view of an illustrative turbomachine 90 thatmay include a turbine rotor blade including integral features accordingto various embodiments of the disclosure. In the example shown,turbomachine 90 includes a gas turbine (GT) system 100 that includes acompressor 102 and a combustor 104. Combustor 104 includes a combustionregion 105 and a fuel nozzle assembly 106. GT system 100 also includes aturbine 108 and a common compressor/turbine shaft (sometimes referred toas a rotor) 110. In one embodiment, GT system 100 is a 7HA or 9HA GTsystem, commercially available from General Electric Company,Greenville, S.C. The present disclosure is not limited to any oneparticular GT system and may be employed in connection with otherengines including, for example, the other HA, F, B, LM, GT, TM andE-class engine models of General Electric Company, and engine models ofother companies. Further, a turbine rotor blade, as described herein,may find application in other forms of turbomachines, e.g., steamturbines, jet engines, compressors, etc.

In operation, air flows through compressor 102 and compressed air issupplied to combustor 104. Specifically, the compressed air is suppliedto fuel nozzle assembly 106 that is integral to combustor 104. Assembly106 is in flow communication with combustion region 105. Fuel nozzleassembly 106 is also in flow communication with a fuel source (not shownin FIG. 1) and channels fuel and air to combustion region 105. Combustor104 ignites and combusts fuel. Combustor 104 is in flow communicationwith turbine 108 for which gas stream thermal energy is converted tomechanical rotational energy. Turbine 108 is rotatably coupled to anddrives rotor 110. Compressor 102 also is rotatably coupled to rotor 110.In the illustrative embodiment, there is a plurality of combustors 104and fuel nozzle assemblies 106.

FIG. 2 shows a cross-sectional view of an illustrative turbine 108 withthree stages that may be used with GT system 100 in FIG. 1. Each stageincludes sets of stationary vanes or nozzles 112 and turbine rotorblades 120. Stationary nozzles 112 may be held in turbine 108 by aradially outer platform 114 and a radially inner platform 116.Stationary nozzles 112 may include one or more circumferentially spacedairfoils 118 (FIG. 4). Turbine rotor blades 120 are coupled to rotor 110and extend between rows of stationary nozzles 112. Combustion gases aredirected by stationary nozzles 112 against turbine rotor blades 120 toturn rotor 110 (FIG. 1).

FIG. 3 shows a perspective view of an illustrative turbine rotor blade120 of GT system 100 in which integral features according to variousembodiments of the disclosure may be employed. FIG. 4 shows an axialcross-sectional view of turbine rotor blade 120 including an integralfeature in the form of an integral impingement cooling structure 160according to various embodiments of the disclosure. Turbine rotor blade120 includes an airfoil body 122 including a concave pressure side outerwall 124 and a convex suction side outer wall 126 that connect alongleading and trailing edges 128, 130. As shown in FIG. 4, outer walls124, 126 have an airfoil inner surface 132 defining a radially extendingchamber 134 for receiving a coolant flow 136. Referring again to FIG. 3,turbine rotor blade 120 may also include a tip end 140 at a radial outerend 142 of airfoil body 122. Turbine rotor blade 120 may also include aturbine rotor blade root 144 (hereinafter “root 144”) by which turbinerotor blade 120 attaches to rotor 110 (FIG. 1), e.g., by a rotor wheel147 (FIG. 21). For purposes of this disclosure, root 144 may include anyportion of turbine rotor blade 120 including, and radially inward of,platform 150. Root 144 may include a blade mount 146 configured formounting in a corresponding slot in the perimeter of a rotor wheel 147(FIG. 21). Blade mount 146 may have any now known or later developedouter configuration for mounting to rotor disk 147 (FIG. 21) such as butnot limited to a dovetail or fir tree arrangement. Turbine rotor blade120, i.e., root 144 thereof, may further include a shank 148 thatextends between blade mount 146 and a platform 150. Platform 150 isdisposed at the junction of airfoil body 122 and shank 148 and defines aportion of the inboard boundary of the flow path through turbine 108(FIGS. 1-2). Shank 148 is thus located at a radial inner end 152 ofairfoil body 122, and blade mount 146 is located radially inward ofshank 148. Platform 150 extends laterally outward relative to shank 148.As will be described further herein, radially extending chamber 134 mayextend at least partially into shank 148 to define a shank inner surface154 (FIG. 5). Outer walls 124 and 126 of airfoil body 122 extend in theradial (Z) direction from platform 150 to tip end 140. It will beappreciated that airfoil body 122 is the active component of turbinerotor blade 120 that intercepts the flow of working fluid and inducesthe rotor to rotate.

B. Integral Impingement Cooling Structure

In certain embodiments, turbine rotor blade 120 may include, inter alia,airfoil body 122 and an integral feature in the form of an integralimpingement cooling structure 160 therein. The impingement coolingstructure is not an insert, but is made integrally through, e.g.,additive manufacture, with the rest of the blade. As noted herein,airfoil body 122 may include concave pressure side outer wall 124 andconvex suction side outer wall 126 that connect along leading andtrailing edges 128, 130. The outer walls 124, 126 have airfoil innersurface 132 defining radially extending chamber 134 for receivingcoolant flow 136. Turbine rotor blade 120 may also include tip end 140at radial outer end 142 of airfoil body 122, and shank 148 at radialinner end 152 of airfoil body 122. Radially extending chamber 134 mayextend at least partially into shank 148 to define shank inner surface154. Integral impingement cooling structure 160 is within radiallyextending chamber 134, and may include hollow body 162 including a firstend 164, a second end 166, an interior surface 168 and an exteriorsurface 170. A plurality of cooling passages 172 extend through hollowbody 162 and are in fluid communication with the radially extendingchamber 134 to allow the coolant flow to pass from interior surface 168of hollow body 162 to impinge on the inner surface of at least airfoilbody 122. In contrast to conventional impingement inserts, first end 164of hollow body 162 is integrally formed to the shank inner surface 154,i.e., via the additive manufacturing. Consequently, exterior surface 170of hollow body 162 may be made to be uniformly spaced from airfoil innersurface 132 between first end 164 and second end 166 of the hollow body162, regardless of the curvature of the airfoil inner surface 132. Inanother embodiment, non-uniform, but custom-built, standoff spacing maybe employed for purposes of providing different impingement cooling,heat pick-up and/or re-use. For example, closer standoff spacing may beemployed where increased impingement cooling is required, and widerstandoff spacing used where less impingement cooling is required.Further, the hollow body 162 may have cooling passages about an entiretyof its periphery and radial span to provide impingement coolingthroughout the blade, not just at a leading edge thereof. Hence, theintegral impingement cooling structure allows for maximum coverage ofimpingement with limited sacrifices normally associated with impingementinserts, and can have a number of variable cooling features. Forexample, the turbine rotor blade may have: variable chordwise width forthe impingement cooling structure or a pin bank aft thereof, tailoredimpingement cooling structure wall thickness for different coolingloads, and varied supports to address different coefficients of thermalexpansion (CTE) between the airfoil body and impingement coolingstructure.

As shown in FIG. 4, and the radial, circumferential cross-sectional viewof FIG. 5, turbine rotor blade 120 may include impingement coolingstructure 160 within radially extending chamber 134. Impingement coolingis typically provided by one or more impingement inserts that areinserted into a radially extending chamber 134 and coupled to airfoilbody 122, e.g., by fasteners or welding. The impingement inserts aretypically linear, but may include some curvature. Where radiallyextending chamber 134 has a curved airfoil inner surface 132 as in FIG.5, it is impossible to have impingement inserts uniformly spaced fromthe inner surface along an entire radial span of the blade. In order toaddress this challenge, impingement cooling structure 160 according toembodiments of the disclosure is integrally formed with the rest ofturbine rotor blade 120, via additive manufacturing.

As shown in FIG. 5, impingement cooling structure 160 includes hollowbody 162 including first end 164, second end 166, interior surface 168and exterior surface 170. Impingement cooling structure 160 alsoincludes plurality of cooling passages 172 through hollow body 162 andin fluid communication with radially extending chamber 134 to allowcoolant flow 136 to pass from interior surface 168 of the hollow body toimpinge on at least airfoil inner surface 132, e.g., it may impingeinner surfaces in, inter alia, airfoil body 122, tip end 140, shank 148,and/or platform 150. In contrast to conventional turbine rotor blades,and as shown in FIG. 4 and especially FIG. 5, exterior surface 170 ofhollow body 162 is uniformly spaced from airfoil inner surface 132between first end 164 and second end 166 of hollow body 162. That is,using additive manufacturing rather than mechanical insertion,impingement cooling structure 160 can be formed (simultaneously with,e.g., airfoil inner surface 132) to have the same curvature, bends,twists and any other shape or dimension, to match that of the innersurface adjacent thereto. Notably, impingement cooling structure 160 canbe uniformly spaced from airfoil inner surface 132 along an entireradial span that it covers, ensuring the desired Z/D parameter over allof turbine rotor blade 120. The Z/D parameter is a ratio of a standoffdistance Z between exterior surface 170 and an inner surface (e.g.,airfoil inner surface 132, shank inner surface 154, etc.) of turbinerotor blade 120 and a diameter D of cooling passages 172 (holes) inimpingement cooling structure 160. In one example, Z/D ranges fromapproximately 1 to approximately 10. In another example, Z/D may rangefrom approximately 2 to approximately 6. A standoff distance Z may besmaller than conventionally available for castings, e.g., less thanapproximately 1.27 millimeters (0.05 inches). Smaller diameter D coolingpassages 172 than conventional castings may also be employed, e.g., as afunction of clogging from debris. Advantageously, cooling passages 172may extend around an entire peripheral extent of hollow body 162 suchthat coolant flow 136 exits hollow body 162 in all directions to provideimpingement cooling to all of airfoil inner surface 132 of airfoil body122. Alternatively, cooling passages 172 may be omitted in areas whereimpingement cooling of inner surfaces 132, 154 is not desired orrequired. Cooling passages 172 may extend along any desired radialextent of hollow body 162.

As shown in FIG. 5, first end 164 of hollow body 162 is integrallyformed to shank inner surface 154. First end 164 meets shank innersurface 154 at a meeting location 174, which extends about the entireperiphery of first end 164, i.e., there are no openings between firstend 164 and shank inner surface 154 at meeting location (other thanperhaps a cooling passage 172). In certain embodiments, first end 164 ofhollow body 162 is integrally formed to shank inner surface 154 radiallyinward of platform 150. However, this particular meeting location 174may not be necessary in all instances, e.g., in some cases, meetinglocation 174 may be radially outward of platform 150. As shown in FIG.5, although not necessary in all cases, second end 166 of hollow body162 may also be integrally formed to inner surface 176 of tip end 140.Cooling passages 172 may optionally provide impingement cooling of tipend 140, or pass coolant to tip end 140 for other forms of cooling.

FIG. 6 shows an enlarged cross-sectional view of meeting location 174(FIG. 5) of impingement cooling structure 160 and shank 148 of turbinerotor blade 120, according to various embodiments of the disclosure. Asshown in FIGS. 5 and 6, first end 164 of hollow body 162 may extendsubstantially in a radial direction (arrow Z) relative to meetinglocation 174 of first end 164 of hollow body 162 and shank inner surface154. As used herein, “substantially in a radial direction” indicatesthat first end 164 extends radially away from rotor 110 (FIG. 1) withsome degree of tolerance, e.g., +/−5°. In contrast, at least a portionof the shank inner surface 154 extends at an angle α relative to radialdirection Z from meeting location 174 of first end 164 of hollow body162 and shank 148 (FIG. 6). In another embodiment, shank inner surface154 is aligned substantially in a radial direction and first end 164 ofhollow body gradually curves or transitions towards meeting location 174to keep angle α, for example, <30°. In a further embodiment, both shankinner surface 154 and first end 164 of hollow body 162 gradually curveor transition towards meeting location 174. Angle α may be any angledesired and within the range of additive manufacture, e.g., <45° fromvertical. To maintain structural integrity, it is desirable that angle αbe as small as possible, e.g., <10°, <20° or <30°. As shown only in FIG.6, in certain embodiments, a support structure 180 may be positionedbetween first end 164 of hollow body 162 and shank inner surface 154,e.g., radially outward of meeting location 174 and radially inward ofplatform 150. In further embodiments, support structures 180 may bepositioned at any location between exterior surface 170 of hollow body162 and airfoil inner surface 132, shank inner surface 154, etc. Instill further embodiments, at least portions of support structure 180include hollow support elements (e.g., lattice) to enable cooling flowdirectly from chamber 134 to outer walls 124 and/or 126 (FIG. 4) ofairfoil body 122. For example, it may be desirable to provide filmcooling to certain areas of airfoil body 122, such as leading and/ortrailing edges 128, 130, directly from radially extending chamber 134.Support structure 180 may include any now known or later developedelement(s) capable of positioning first end 164 of hollow body 162relative to shank inner surface 154. Support structure 180 may include,but is not limited to: a lattice structure, straight or arced bar(s),etc. Support structure 180 may also be integrally formed via additivemanufacture.

Impingement cooling structure 160 may also include a variety of optionalalternative integral cooling features. In one example, impingementcooling structure 160 may be optionally formed with varying wallthicknesses. Varying wall thicknesses may be advantageous to, forexample, accommodate varying CTEs between impingement cooling structure160 and hotter airfoil body 122, shank 148 and/or platform 150. Asobserved in FIG. 5, airfoil body 122, shank 148 and/or platform 150 mayhave a wide variety of wall thicknesses, and may have varyingthicknesses over an extent thereof. FIG. 7 shows an enlarged, partialcross-sectional view of a portion of impingement cooling structure 160adjacent airfoil body 122, platform 150 or shank 148. As noted, incertain embodiments, as shown in FIG. 7, hollow body 162 may include atleast one first portion 182 having a first wall thickness W1 betweeninterior surface 168 and exterior surface 170 thereof, and at least onesecond portion 184 having a second wall thickness W2 between interiorsurface 168 and exterior surface 170 thereof. In the example shown,first wall thickness W1 is greater than second wall thickness W2. Anynumber of thicker and/or thinner portions 182, 184 may be provided inimpingement cooling structure 160. The thickness of portions 182, 184may be any dimension desired to address the structural and/or thermalrequirements of the location.

In another example optional structure, additional support may be desiredand/or required to support integral impingement cooling structure 160relative to inner surfaces 132, 154. For example, additional support maybe desired and/or required at thinner wall thickness portions 184 (FIG.7) of impingement cooling structure 160. To this end, as shown in FIG.7, turbine rotor blade 120 may also include a support 186 on exteriorsurface 170 of hollow body 162 in at least one portion 184 havingthinner wall thickness W2. Support 186 may be integrally formed withhollow body 162 (and rest of turbine rotor blade 120) to space exteriorsurface 170 of hollow body 162 from, e.g., airfoil inner surface 132,between first end 164 (FIG. 5) and second end 166 of hollow body 162.Any number of supports 186 may be provided in thinner wall portion(s)184. Support(s) 186 may include a passage(s) 188 therethrough in fluidcommunication with one of the plurality of cooling passages 172, i.e.,to allow coolant flow 136 to pass therethrough and impinge innersurface(s) 132, 154. In certain embodiments, regardless of wallthickness, turbine rotor blade 120 may include a support(s) 189 onexterior surface 170 of hollow body 162. Support 189 may be integrallyformed with hollow body 162 (and rest of turbine rotor blade 120) tospace exterior surface 170 of hollow body 162 from, e.g., airfoil innersurface 132, between first end 164 (FIG. 5) and second end 166 of hollowbody 162. Supports 186, 189 may take any form that allows: reduction ofstress between hotter outer walls 124, 126 of airfoil body 122 andcooler impingement cooling structure 160, provide any necessary thermalexpansion, provide structural support, and/or desired spacing of hollowbody 162 from inner surface(s) 132, 154. Supports 186, 189 can have anydesired dimension and/or shape such as but not limited to: tubes, bars,etc.

FIG. 8 shows an enlarged cross-sectional view of another alternativeembodiment including a reinforcement member 190 surrounding at least oneof cooling passages 172. Reinforcement member 190 may include anystructurally strengthening member such as a thicker wall, etc. Certainembodiments, as shown in FIG. 4, may also include a stiffener rib 192integrally formed to interior surface 168 of hollow body 162. Any numberof stiffener ribs 192 may be provided, and each may extend any desiredradial extent of hollow body 162. Supports 186, 189, reinforcementmember 190 and/or stiffener rib 192 may be integrally formed with therest of turbine rotor blade 120 via additive manufacture.

FIG. 9 shows a cross-sectional view of turbine rotor blade 120 includingintegral impingement cooling structure 160 and additional optionalalternative integral cooling features. In one alternative embodiment,impingement cooling structure 160 may be optionally formed with varyingspacing Z from inner surface 132, 154. The spacing Z may be customizedto provide the desired Z/D parameter and desired cooling at variouslocations. For example, turbine rotor blade 120 may have a number ofhigh heat load regions 195, i.e., regions that experience highertemperatures and require more cooling compared to other regions of theblade. In the example shown, high heat load regions 195 include regions:near leading edge 128 (195A), pressure side outer wall 124 near trailingedge 130 (195B), and suction side wall 126 downstream of leading edge128 (195C). At high heat load regions 195, a first spacing Z1 may beemployed between integral impingement cooling structure 160 and innersurface 132, 154 at high heat load regions 195, while a second, largerspacing Z2 is used at other locations that do not have such a high heatload. In this manner, more cooling can be provided where necessary,i.e., at high heat load regions 195, using first spacing Z1 with thespacing increasing between impingement cooling structure 160 and innersurface 132, 154 to second, larger spacing Z2 for lower heat loadregions. As shown in FIG. 9, the larger second spacing Z2 may allowcoolant flow 136 to limit or reduce heat absorption as it movesdownstream toward trailing edge 130, allowing coolant flow 136 to becooler and have more heat absorbing capacity for downstream regions,e.g., serpentine cooling passage 200 and/or pin bank 206 (describedherein). A transition between spacings Z1 and Z2 can be at any desiredrate, e.g., gradual over a relatively long distance, abrupt at aparticular location, or at any rate therebetween. Second spacing Z2 maybe anywhere from, for example, 1.01 to 3.00 times first spacing Z1. TheZ/D parameter can be customized for each region of concern. As noted, inone example, Z/D ranges from approximately 1 to approximately 10. Inanother example, Z/D may range from approximately 2 to approximately 6.Diameter D of cooling passages 172 can also be configured to customizethe Z/D parameter for different regions.

FIG. 9 also shows turbine rotor blade 120 including one or morepost-impingement target features 196 on inner surface 132.Post-impingement target features 196 may include any now known or laterdeveloped structure on inner surface 132 to promote cooling. In theexample shown, impingement target features 192 include bumps, but theycould include any structure. In one scenario, hollow body 162 mayinclude a local bulge 198 to match impingement target feature(s) 196contour, and thus maintain spacing Z (i.e., Z1 as shown). While twotarget features 196 and bulge 198 pairs are shown, any number may beemployed. In one embodiment, post-impingement target features 196 mayalso optionally include additional integral cooling features, such asbut not limited to film cooling holes 199. Film cooling holes 199 directcoolant flow 136, post-impingement with post-impingement target features196, i.e., inner surface 132 thereof, to create a cooling film 201 oversidewall(s) 124, 126. Any number of film cooling holes 199 may beapplied within each post-impingement cooling features 196.

FIG. 10 shows a first radial, cross-sectional view along view line 10-10in FIG. 4, and FIG. 11 shows a second, radial cross-sectional view alongview line 11-11 in FIG. 4, the latter of which is in a slightlydifferent plane than FIG. 10 and in the opposite direction. As shown inFIG. 10, in certain embodiments, hollow body 162 has a chordwise widthWC1 that is smaller near tip end 140 than shank 148. Most conventionalimpingement inserts have an opposite chordwise width arrangement toallow them to be inserted through an open tip end of the airfoil body.Further, hollow body 162 may have alternating wider and narrow chordwisewidths WC1 over its radial span (up and down page in FIGS. 10-11).Consequently, an axially aft end 194 of hollow body 162 may vary in achordwise location along a radial span of hollow body 162. In thismanner, impingement cooling structure 160 can have a shape (chordwisewidth WC1) that curves over its radial span to be uniformly spaced fromairfoil inner surface 132 and/or shank inner surface 154 (FIG. 5),regardless of the latter's shape. Conventional impingement insertscannot provide such features.

As shown in FIGS. 4, 10 and 11, airfoil body 122 further includes atleast one chordwise extending, serpentine cooling passage 200 extendingfrom airfoil inner surface 132 aft of hollow body 162 toward trailingedge 130. As shown best in FIGS. 10 and 11, each chordwise extending,serpentine cooling passage 200 may have a same chordwise width WC2,e.g., shorter than chordwise width WC1 of hollow body 162. Airfoil body122 may also include a plurality of radially spaced, trailing edgecooling passages 202 extending through trailing edge 130, i.e., fromserpentine cooling passage(s) 200. Each of the plurality of trailingedge cooling passages 202 has a same chordwise width WC3, i.e., along aradial span of turbine rotor blade 120. As illustrated in FIGS. 10 and11, a space 204 between trailing edge cooling passages 202 andserpentine cooling passage(s) 200 has a varying chordwise width WC4,i.e., along a radial span of turbine rotor blade 120. Turbine rotorblade 120 may further include a pin bank 206 between a forward end ofplurality of trailing edge cooling passages 202 (right side thereof inFIG. 10, left side thereof in FIG. 11) and an aft end of chordwiseextending, serpentine cooling passage(s) 200 (left side thereof in FIG.10, right side thereof in FIG. 11). Consequently, as illustrated, inFIGS. 10-11, pin bank 206 may have the varying chordwise width WC4 alonga radial span thereof. FIGS. 10 and 11 illustrate a variety of coolingfeatures including but not limited to: film cooling via openings 208,main chord impingement cooling via impingement cooling structure 160,near trailing edge 130 cooling via serpentine cooling passage(s) 200,and trailing edge 130 pin bank cooling via pin bank 206.

Additive manufacturing (AM) includes a wide variety of processes ofproducing a component through the successive layering of material ratherthan the removal of material. As such, additive manufacturing can createcomplex geometries, such as those described herein relative to turbinerotor blade 120, without the use of any sort of tools, molds orfixtures, and with little or no waste material. Instead of machiningcomponents from solid billets of material, much of which is cut away anddiscarded, the only material used in additive manufacturing is what isrequired to shape the component. Additive manufacturing techniquestypically include taking a three-dimensional computer aided design (CAD)file of the component (e.g., turbine rotor blade 120) to be formed,electronically slicing the component into layers, e.g., 18-102micrometers thick, and creating a file with a two-dimensional image ofeach layer, including vectors, images or coordinates. The file may thenbe loaded into a preparation software system that interprets the filesuch that the component can be built by different types of additivemanufacturing systems. In 3D printing, rapid prototyping (RP), anddirect digital manufacturing (DDM) forms of additive manufacturing,material layers are selectively dispensed, sintered, formed, deposited,etc., to create the component. While other manufacturing processes suchas casting may also be employed, turbine rotor blade 120 may beadvantageously made by additive manufacturing.

In metal powder additive manufacturing techniques, such as direct metallaser melting (DMLM) (also referred to as selective laser melting(SLM)), direct metal laser sintering (DMLS), selective laser sintering(SLS), electron beam melting (EBM), and perhaps other forms of additivemanufacturing, metal powder layers are sequentially melted together toform the component. More specifically, fine metal powder layers aresequentially melted after being uniformly distributed using anapplicator on a metal powder bed. Each applicator includes an applicatorelement in the form of a lip, brush, blade or roller made of metal,plastic, ceramic, carbon fibers or rubber that spreads the metal powderevenly over the build platform. The metal powder bed can be moved in avertical axis. The process takes place in a processing chamber having aprecisely controlled atmosphere. Once each layer is created, each twodimensional slice of the component geometry can be fused by selectivelymelting the metal powder. The melting may be performed by a high poweredmelting beam, such as a 100 Watt ytterbium laser, to fully weld (melt)the metal powder to form a solid metal. The melting beam moves in theX-Y direction using scanning mirrors, and has an intensity sufficient tofully weld (melt) the metal powder to form a solid metal. The metalpowder bed may be lowered for each subsequent two dimensional layer, andthe process repeats until the component is completely formed. In orderto create certain larger blades faster, some metal additivemanufacturing systems employ a pair of high powered lasers that worktogether to form a blade. Here, a method of making turbine rotor blade120 may include sequentially creating a layer of material and applying aheat source to sinter the layer of materials to form the structuredescribed herein. Thus, additive manufacturing results in airfoil body122, tip end 140, shank 148 and impingement cooling structure 160including a plurality of integral material layers.

Turbine rotor blade 120 may be made of a metal which may include a puremetal or an alloy, capable of withstanding the environment in whichemployed. In one example, the metal may include practically anynon-reactive metal powder, i.e., non-explosive or non-conductive powder,such as but not limited to: a cobalt chromium molybdenum (CoCrMo) alloy,stainless steel, an austenite nickel-chromium based alloy such as anickel-chromium-molybdenum-niobium alloy (NiCrMoNb) (e.g., Inconel 625or Inconel 718), a nickel-chromium-iron-molybdenum alloy (NiCrFeMo)(e.g., Hastelloy® X available from Haynes International, Inc.), or anickel-chromium-cobalt-molybdenum alloy (NiCrCoMo) (e.g., Haynes 282available from Haynes International, Inc.), etc. In another example, themetal may include practically any metal such as but not limited to: toolsteel (e.g., H13), titanium alloy (e.g., Ti₆Al₄V), stainless steel(e.g., 316L) cobalt-chrome alloy (e.g., CoCrMo), and aluminum alloy(e.g., AlSi₁₀Mg).

In contrast to conventional impingement inserts, first end 164 of hollowbody 162 is integrally formed to shank inner surface 154, i.e., viaadditive manufacturing. Consequently, exterior surface 170 of hollowbody 162 may be made to be uniformly spaced from inner surface(s)between first end 164 and second end 166 of hollow body 162, regardlessof the curvature of, for example, airfoil inner surface 132 and/or shankinner surface 154. Further, hollow body 162 may have cooling passages172 about an entirety of its periphery and radial span to provideimpingement cooling throughout the blade, not just at a leading edgethereof. Hence, the integral impingement cooling structure 160 allowsfor maximum coverage of impingement with no sacrifices normallyassociated with impingement inserts, and can have a number of variablecooling features. For example, turbine rotor blade 120 may have: avariable chordwise width for the impingement cooling structure 160(i.e., width WC1) or a pin bank 206 (i.e., WC4) aft of structure 160, atailored impingement cooling structure 160, different wall thicknessesfor different cooling and/or structural loads, and varied supports 186,189 to address different coefficients of thermal expansion (CTE) betweenairfoil body 122 and impingement cooling structure 160. Additionalcooling features, such as turbulators (not shown), may also be providedand customized around each cooling passage 172 to optimize impingementcooling. Turbine rotor blade 120 also may include axial venting throughtrailing edge 130, as described relative to FIGS. 10-11.

C. Platform with Cooling Passages Having Non-Linear Configuration

Referring to FIGS. 12-18, another integral feature according toembodiments of the disclosure is illustrated. Similar to the previousembodiment, turbine rotor blade 120 may include airfoil body 122 withradially extending chamber 134 for receiving coolant flow 136. As shownbest in FIGS. 3 and 10-13, platform 150 extends laterally outwardrelative to airfoil body 122 and terminates at at least one slash face230 (e.g., FIGS. 11, 12). FIG. 12 shows a perspective, transparent viewof a pressure side 232 of platform 150, and FIG. 13 shows a top down,transparent view of a suction side 231 of platform 150. As shown inFIGS. 12 and 13, a cooling circuit 234 is located within platform 150and is in fluid communication with a source of coolant 236. Source ofcoolant 236 may take any of a variety of forms. In one example, whereturbine rotor blade 120 includes impingement cooling structure 160 inradially extending chamber 134, source of coolant 236 to cooling circuit234 may provide the coolant after passing through impingement coolingstructure 160, i.e., the coolant is post-impingement coolant. In anotherembodiment, source of coolant 236 may be radially extending chamber 134.For example, where impingement cooling structure 160 is not provided, orit is provided radially outward of platform 150, source of the coolant236 to cooling circuit 234 may provide the coolant directly fromradially extending chamber 134. Other sources of coolant 236 may also beused, e.g., wheel space portion between shanks 148 of adjacent turbinerotor blades 120. Cooling circuit 234 may take any now known or laterdeveloped form. In the example shown in FIG. 13, cooling circuit 234includes a sinusoidal path through platform 150. In contrast, in FIG.12, cooling circuit 234 includes an elbow path. Cooling circuit 234 mayhave a less complex path or a more complex path, and may extend wherenecessary to cool platform 150.

Turbine rotor blade 120 also includes cooling passage(s) 240 fromcooling circuit 234 through a surface 242 of slash face(s) 230, i.e., tocool slash faces 230 and other structure. Cooling passage(s) 240 are inplatform 150 and in fluid communication with cooling circuit 234. Incontrast to conventional, linear cooling passages, cooling passage(s)240 extend in a non-linear configuration from cooling circuit 234 toexit through at least one slash face 230 of platform 150, providingimproved cooling compared to linear cooling passages. For example, inFIG. 12, cooling passage(s) 240 have a (gently) curved shaped. Anynumber of cooling passage(s) 240 may be employed, to provide the desiredcooling. Further, they may have any uniform or non-uniformcross-sectional shape desired, and may be uniformly or non-uniformlyspaced, to provide the desired cooling. The non-linear configuration ismade possible by, for example, additive manufacturing. As noted, airfoilbody 122 and platform 150, including the parts that define coolingpassages(s) 240, may include a plurality of integral material layers.

FIG. 14 shows an enlarged cross-sectional view of a slash face 230. Insome embodiments, as shown in FIG. 14, slash face(s) 230 may include anextension member 244. Extension member 244 may define a damper pin seat246 configured to receive a damper pin 248 (shown in phantom) that sealswith the damper pin seat of an adjacent turbine rotor blade (not shown).Where provided, cooling passage(s) 240 may extend through extensionmember 244. In this regard, cooling passage(s) 240 may have a non-linearconfiguration that extends from cooling circuit 234 radially outward andabout damper pin seat 246 to outer surface 242 of slash face 230, i.e.,in a curved shape that is sharper or more turning than FIG. 12.

Cooling passage(s) 240 may take any of a number of non-linearconfigurations to provide the desired cooling. The non-linearconfiguration, e.g., curved shape, may extend in any desired directionwithin platform 150, e.g., radially (inward or outward), axially (aft orforward) or circumferentially (clockwise or counterclockwise), or acombination of the directions. Cooling passage(s) 240 may all have thesame shape to provide the same cooling attributes at each location whereprovided, or they may vary in shape within platform 150 to providecustom cooling for each location where they are provided. In addition tothe curved shape shown in FIGS. 12 and 13, in another embodiment shownin FIG. 15, cooling passage(s) 240 may have a helical (corkscrew) shape,i.e., with a number of helical coils 250. Any number of helical coils250 may be used for each cooling passage 240. As shown in FIGS. 13 and16, cooling passage(s) 240 may have at least one first turn 252 (FIG.16) in a first direction FD, and at least one second turn 254 (FIG. 16)in a second, opposite direction SD, creating a generally zig-zag path.Any number of first and second turns 252, 254 (FIG. 16) may be used foreach cooling passage 240. As shown on the left side of FIG. 16, anamplitude A of each turn 252, 254 can be consistent so as to form asinusoidal shape with the at least one first and second turns of equalamplitude A. Alternatively, as shown on the right side of FIG. 16, theamplitude of each turn 252, 254 can be inconsistent so as to form a morerandom zig-zag path with turns 252, 254. As also shown in the right sideof FIG. 16, an input 260 and an exit 262 of each cooling passage 240need not be aligned. In another embodiment, shown in FIG. 17, coolingpassage(s) 240 may have a plurality of branches 264, e.g., like a tree.Any branching configuration may be employed.

FIG. 18 shows an embodiment in which cooling passage(s) 240 have acurved shape, e.g., more planar within platform 150 than in FIG. 12.FIG. 18 also illustrates exit 262 of cooling passage(s) 240 may meetsslash face 230 of platform 150 at an angle α that is not 90°. Inembodiment, angle α is less than 15°. Angle α may be customized toprovide the desired cooling to platform 150 and/or film cooling to slashface 230. While shown separately, any of the cooling passage examples oraspects thereof may be combined with the other examples.

Cooling circuit 234 and cooling passage(s) 240 may be provided inpressure side 232 of platform alone (FIG. 13 alone), in suction side 231of platform 150 alone (FIG. 12 alone), or in both sides 231, 232 ofplatform 150 (FIGS. 12-13). If provided on only one side of platform150, any other conventional structure may be provided in the other sideof the platform. In the latter case, as shown collectively in FIGS. 12and 13, cooling circuit 234 may include a first portion 234SS in suctionside 231 of platform 150 and a second portion 234PS on pressure side 232of platform 150. Portions 234SS and 234PS may be separated or fluidlycoupled. In this regard, slash face 230 includes a suction side slashface 230SS and a pressure side slash face 230PS. Here, coolingpassage(s) 240 in platform 150 may include: at least one first coolingpassage 240 in fluid communication with first portion 234SS of thecooling circuit and exiting suction side slash face 230SS, and at leastone second cooling passage 240 in fluid communication with secondportion 234PS of cooling circuit 234 and exiting pressure side slashface 230PS.

Non-linear cooling passages 240 allow coolant to be directed whereneeded in platform 150, in contrast to conventional, drilled linearcoolant passages. The additive manufacture of coolant passages 140 allowthem to have a wide variety of non-linear configurations that directcooling where necessary and provide enhanced cooling through theirshape.

D. Angel Wing with Coolant Transfer Passage

Referring to FIGS. 2 and 19-22, another integral feature according toembodiments of the disclosure includes an angel wing 280 having acoolant transfer passage therein. FIG. 19 shows a radial cross-sectionthrough turbine rotor blade 120 including an angel wing 280, FIG. 20shows a transparent perspective view of turbine rotor blade 120including angel wing 280, FIG. 21 shows an axial view of a set ofturbine rotor blades 120A-C including angel wing(s) 280, and FIG. 22shows a top down view of a turbine rotor blade 120 including angel wings280. With further respect to FIG. 21, a set of turbine rotor bladesincludes: a first turbine rotor blade 120A, a second turbine rotor blade120B and a third turbine rotor blade 120C (collectively or individually,turbine rotor blade 120). First turbine rotor blade 120A is positionedbetween second and third turbine rotor blades 120B, 120C. In thisembodiment, as shown in FIGS. 19-20, each turbine rotor blade 120 mayinclude airfoil body 122 including concave pressure side outer wall 124(FIG. 3) and convex suction side outer wall 126 (FIG. 3) that connectalong leading and trailing edges 128, 130 (FIG. 3). Turbine rotor blade120 may also include shank 148 at radial inner end 152 of airfoil body122. In addition, turbine rotor blade 120 includes at least one angelwing 280 extending laterally from at least one side 282, 284 of shank148.

As shown for one nozzle-blade interface in FIG. 2, an opening 286 existsat the interface between adjacent nozzles 112 and turbine rotor blades120 that can allow hot working fluid to exit the hot gas path and entera wheel space 300 of turbine 108. In order to limit this leakage of hotgas, turbine rotor blade 120 typically includes axially projecting angelwing seals 280, also simply referred to as ‘angel wings’. Angel wings280 cooperate with projecting segments or ‘discouragers’ 288 whichextend from nozzle 112. Angel wings 280 and discouragers 288 overlap (ornearly overlap), but do not touch each other, thus restricting fluidflow.

Turning to FIGS. 19-22, in accordance with embodiments of thedisclosure, turbine rotor blade 120 may also include a coolant transferpassage 290 defined through the at least one angel wing 280. Coolanttransfer passage 290, e.g., for a first turbine rotor blade 120A (FIG.21), fluidly couples a first wheel space portion 292 defined betweenshank 148A (FIG. 21) and a first adjacent shank 148B (FIG. 21) of afirst adjacent turbine rotor blade 120B and a second wheel space portion294 defined between shank 148A (FIG. 21) and a second adjacent shank148C of a second adjacent turbine rotor blade 120C (FIG. 21). As shownbest by observing FIGS. 2, 21 and 22, each wheel space portion 292, 294is part of wheel space 300. Wheel space 300 is defined:circumferentially between shanks 148A-C (FIG. 21) of adjacent turbinerotor blades 120A-C, axially between shank 148 and an adjacent nozzle112, and radially by platform 150 and rotor disks 147. Wheel spaceportions 292, 294 are that part of wheel space 300 that is axiallybeside a particular blade's shank 148.

As shown best in FIG. 22, coolant transfer passage 290 includes a firstopen end 310 in fluid communication with first wheel space portion 292and a second open end 312 in fluid communication with second wheel spaceportion 294. Thus, coolant transfer passage 290 allows a wheel spacecoolant 316 to pass between wheel space portions 292, 294 oncircumferentially opposing sides of shank 148. First open end 310 andsecond open end 312 may face in an axial-circumferential directionrelative to airfoil body 122, or any direction that will allow wheelspace coolant 316 to pass between wheel space portions 292, 294. Wheelspace coolant 316 may be any now known or later developed coolant, e.g.,routed from compressor 102 (FIG. 1). As noted previously, outer walls124, 126 of airfoil body 122 define radially extending chamber 134 thatmay extend into shank 148. As shown in FIG. 22, coolant transfer passage290 is fluidly isolated from radially extending chamber 134, i.e.,coolant flow 136 (FIG. 4) from chamber 134 does not mix with wheel spacecoolant 316.

Any number of angel wings 280 may be employed. In one example, shown inFIGS. 19, 20 and 22, a first angel wing 280 extends laterally from firstside 282 of shank 148, and a second angel wing 280 extends laterallyfrom a second, opposing side 284 of shank 148. In another example, shownin FIG. 23, a first pair of radially spaced angel wings 280 may extendlaterally from first side 282 of shank 148, and none extend from side284 of shank 148. In another embodiment, as shown in FIG. 24, a firstpair of radially spaced angel wings 280 may extend laterally from firstside 282 of shank 148, and a second pair of radially spaced angel wings280 may extend laterally from a second, opposing side 284 of shank 148.In any event, each angel wing 280 may include a respective coolanttransfer passage 290. Alternatively, while each angel wing 280 is shownincluding a coolant transfer passage 290, selective angel wings may notinclude a coolant transfer passage.

Coolant transfer passage 290 allows wheel space coolant 316 (FIG. 22) tomove between wheel space portions 292, 294, allows cooling of angelwings 280, and reduces the weight turbine rotor blade 120.

E. Hollow Blade Mount with Lattice Support Structure

Referring to FIGS. 19, 20, 25 and 26, another integral feature accordingto embodiments of the disclosure includes a hollow blade mount 146. Inthis embodiment, root 144 is provided including shank 148 havingradially extending chamber 134 defined therein. Blade mount 146 is at aradial inner end of shank 148. In contrast to many conventional blademounts, blade mount 146 has a hollow interior 330 defined therein, e.g.,by inner wall surfaces 344 of blade mount 146. Hollow interior 330 is influid communication with radially extending chamber 134. Hollow interior330 may have any desired interior shape, e.g., expanding radially asshown in FIG. 25. Blade mount 146 may have any now known or laterdeveloped exterior shape configured for mounting to a rotor wheel 147(FIG. 21) coupled to rotor 110 (FIG. 2), e.g., a dovetail, or fir treeshape.

Turbine rotor blade root 144 may further include a lattice supportstructure 340 disposed within hollow interior 330 of the blade mount146. Lattice support structure 340 may take a variety of hollow supportstructure forms. In one example, lattice support structure 340 mayinclude a plurality of radially extending V-shaped sections 342.V-shaped sections 342 may be integral with inner wall surfaces 344 ofblade mount 146. Root 144 including shank 148 and blade mount 146,including lattice support structure 340, may be made by additivemanufacture. Shank 148 and blade mount 146 thus may include a pluralityof integral material layers.

Root 144 according to this embodiment, i.e., with lattice supportstructure 340, may also include platform 150, as described hereinrelative to FIGS. 12-18. Platform 150, as noted, is positioned radiallyoutward of shank 148 and extends laterally outward relative to theshank, terminating at at least one slash face 230. Platform 150 mayinclude cooling circuit 234 defined within the platform and in fluidcommunication with a source of a coolant flow, e.g., radially extendingchamber 134. Cooling passage(s) 240 (FIGS. 12-18) may be defined inplatform 150 and in fluid communication with cooling circuit 234. Asnoted, cooling passage(s) 240 extending in a non-linear configurationfrom cooling circuit 234 to exit through slash face(s) 230 of theplatform. Slash face(s) 230 may include extension member 244 throughwhich cooling passage(s) 240 extend. Cooling passage(s) 240 may have: ahelical shape (FIG. 15); at least one first turn in a first direction,and at least one second turn in a second, opposite direction (FIG. 16);a plurality of branches (FIG. 17); or a curved shape (e.g., FIGS. 12,14, 18).

Root 144 according to this embodiment, i.e., with lattice supportstructure 340, may also include angel wing(s) 280 extending laterallyfrom at least one side of shank 148, as described herein relative toFIGS. 19-24. As noted, a coolant transfer passage 290 may be definedthrough angel wing(s) 280. As shown in FIG. 21, coolant transfer passage290 fluidly couples a first wheel space portion 292 defined betweenshank 148A and a first adjacent shank 148B of a first adjacent turbinerotor blade root 144B and a second wheel space portion 294 definedbetween shank 148A and a second adjacent shank 148C of a second adjacentturbine rotor blade root 144C. Coolant transfer passage 290 includes afirst open end 310 in fluid communication with first wheel space portion292 and a second open end 312 in fluid communication with second wheelspace portion 294. As shown in FIG. 22, first open end 310 and secondopen end 312 may face in a circumferential direction relative to shank148. Coolant transfer passage 290 may be fluidly isolated from radiallyextending chamber 134 in shank 148.

Root 144 according to this embodiment, i.e., with lattice supportstructure 340, may also include both platform 150 and angel wing(s) 280,as described herein. Additive manufacture allows for formation of root144 with shank 148, hollow blade mount 146, lattice support structure340, and platform 150 and/or angel wing(s) 280, creating a plurality ofintegral material layers for whatever features are provided.

Root 144 including integral lattice support structure 340 in hollowinterior 330 of blade mount 146 provides a lighter turbine rotor blade120, and additional cooling of blade mount 146.

While the various embodiments have been described and illustrated hereinas used together, it is understood that the various embodiments can beused alone or in a combination.

Approximating language, as used herein throughout the specification andclaims, may be applied to modify any quantitative representation thatcould permissibly vary without resulting in a change in the basicfunction to which it is related. Accordingly, a value modified by a termor terms, such as “about,” “approximately” and “substantially,” are notto be limited to the precise value specified. In at least someinstances, the approximating language may correspond to the precision ofan instrument for measuring the value. Here and throughout thespecification and claims, range limitations may be combined and/orinterchanged; such ranges are identified and include all the sub-rangescontained therein unless context or language indicates otherwise.“Approximately” as applied to a particular value of a range applies toboth end values, and unless otherwise dependent on the precision of theinstrument measuring the value, may indicate +/−10% of the statedvalue(s).

The corresponding structures, materials, acts, and equivalents of allmeans or step plus function elements in the claims below are intended toinclude any structure, material, or act for performing the function incombination with other claimed elements as specifically claimed. Thedescription of the present disclosure has been presented for purposes ofillustration and description but is not intended to be exhaustive orlimited to the disclosure in the form disclosed. Many modifications andvariations will be apparent to those of ordinary skill in the artwithout departing from the scope and spirit of the disclosure. Theembodiment was chosen and described in order to best explain theprinciples of the disclosure and the practical application, and toenable others of ordinary skill in the art to understand the disclosurefor various embodiments with various modifications as are suited to theparticular use contemplated.

What is claimed is:
 1. A turbine rotor blade, comprising: an airfoilbody including a concave pressure side outer wall and a convex suctionside outer wall that connect along leading and trailing edges, theconcave pressure side and the convex suction side outer walls having anairfoil inner surface defining a radially extending chamber forreceiving a coolant flow; wherein the airfoil body further includes atleast one chordwise extending, serpentine cooling passage; a pluralityof radially spaced, trailing edge cooling passages extending through thetrailing edge; a space between the plurality of trailing edge coolingpassages and the at least one chordwise extending, serpentine coolingpassage, wherein the space has a varying chordwise width along a radialspan thereof; a tip end at a radial outer end of the airfoil body; ashank at a radial inner end of the airfoil body, the radially extendingchamber extending at least partially into the shank to define a shankinner surface; and an impingement cooling structure within the radiallyextending chamber, the impingement cooling structure including: a hollowbody including a first end, a second end, an interior surface and anexterior surface, and a plurality of cooling passages through the hollowbody and in fluid communication with the radially extending chamber toallow the coolant flow to pass from the interior surface of the hollowbody to impinge on at least the airfoil inner surface, wherein the firstend of the hollow body is integrally formed to the shank inner surface,wherein a first chordwise width of the hollow body at the tip end isless than a second chordwise width of the hollow body at the shank, andwherein the exterior surface of the hollow body is uniformly spaced fromthe airfoil inner surface between the first end and the second end ofthe hollow body.
 2. The turbine rotor blade of claim 1, wherein thesecond end of the hollow body is integrally formed to an inner surfaceof the tip end.
 3. The turbine rotor blade of claim 1, wherein the firstend of the hollow body extends substantially in a radial directionrelative to a meeting location of the first end of the hollow body andthe shank inner surface.
 4. The turbine rotor blade of claim 3, whereinat least a portion of the shank inner surface extends at an anglerelative to the radial direction from the meeting location of the firstend of the hollow body and the shank.
 5. The turbine rotor blade ofclaim 3, further comprising: a platform between the shank and the radialinner end of the airfoil body, the platform extending laterally outwardrelative to the shank, and a support structure between the first end ofthe hollow body and the shank inner surface, the support structurepositioned radially outward of the meeting location and radially inwardof the platform.
 6. The turbine rotor blade of claim 1, furthercomprising a platform between the shank and the radial inner end of theairfoil body, the platform extending laterally outward relative to theshank, wherein the first end of the hollow body is integrally formed tothe shank inner surface radially inward of the platform.
 7. The turbinerotor blade of claim 1, wherein the hollow body includes at least onefirst portion having a first wall thickness between the interior surfaceand the exterior surface thereof, and at least one second portion havinga second wall thickness between the interior surface and the exteriorsurface thereof, the first wall thickness being greater than the secondwall thickness.
 8. The turbine rotor blade of claim 7, furthercomprising a support on the exterior surface of the hollow body in theat least one second portion having the second wall thickness, thesupport integrally formed with the hollow body to space the exteriorsurface of the hollow body from the airfoil inner surface between thefirst end and the second end of the hollow body, wherein the supportincludes a passage therethrough in fluid communication with one of theplurality of cooling passages.
 9. The turbine rotor blade of claim 1,further comprising a support on the exterior surface of the hollow body,the support integrally formed with the hollow body to space the exteriorsurface of the hollow body from the airfoil inner surface between thefirst end and the second end of the hollow body.
 10. The turbine rotorblade of claim 1, further comprising a reinforcement member surroundingat least one of the plurality of cooling passages.
 11. The turbine rotorblade of claim 1, wherein the airfoil body, the tip end, the shank andthe impingement cooling structure include a plurality of integralmaterial layers.
 12. The turbine rotor blade of claim 1, furthercomprising a stiffener rib integrally formed to the interior surface ofthe hollow body.
 13. The turbine rotor blade of claim 1, wherein theplurality of cooling passages through the hollow body and in fluidcommunication with the radially extending chamber extend around anentire peripheral extent of the hollow body.
 14. The turbine rotor bladeof claim 1, wherein an axially aft end trailing edge portion of thehollow body varies in a chordwise location along a radial span of thehollow body, and wherein each chordwise extending, serpentine coolingpassage of the at least one chordwise extending, serpentine coolingpassages has a same chordwise width.
 15. The turbine rotor blade ofclaim 14, wherein each of the plurality of trailing edge coolingpassages has a same chordwise width.
 16. The turbine rotor blade ofclaim 15, wherein the airfoil body further includes a pin bank between aforward end of the plurality of trailing edge cooling passages and anaft end of the at least one chordwise extending, serpentine coolingpassage, wherein the pin bank has a varying chordwise width along aradial span thereof.
 17. An additively manufactured turbine rotor blade,comprising: an airfoil body including: a concave pressure side outerwall and a convex suction side outer wall that connect along leading andtrailing edges, the concave pressure side and the convex suction sideouter walls having an airfoil inner surface defining a radiallyextending chamber for receiving a coolant flow; at least one chordwiseextending, serpentine cooling passage; a plurality of radially spaced,trailing edge cooling passages extending through the trailing edge,wherein each of the plurality of trailing edge cooling passages has asame chordwise width; a space between the plurality of trailing edgecooling passages and the at least one chordwise extending, serpentinecooling passage, wherein the space has a varying chordwise width along aradial span thereof; a pin bank between a forward end of the pluralityof trailing edge cooling passages and an aft end of the at least onechordwise extending, serpentine cooling passage, wherein the pin bankhas a varying chordwise width along a radial span thereof; and anintegral impingement cooling structure within the radially extendingchamber, the integral impingement cooling structure including: a hollowbody including a first end, a second end, an interior surface and anexterior surface, and a plurality of cooling passages through the hollowbody and in fluid communication with the radially extending chamber toallow the coolant flow to pass from the interior surface of the hollowbody to impinge on at least the airfoil inner surface, wherein theexterior surface of the hollow body is uniformly spaced from the airfoilinner surface between the first end and the second end of the hollowbody.
 18. The additively manufactured turbine rotor blade of claim 17,wherein the plurality of cooling passages through the hollow body are influid communication with the radially extending chamber extend around anentire peripheral extent of the hollow body.
 19. The additivelymanufactured turbine rotor blade of claim 17, wherein the second end ofthe hollow body is integrally formed to an inner surface of the tip endof the airfoil body.
 20. The additively manufactured turbine rotor bladeof claim 17, further comprising a tip end at a radial outer end of theairfoil body and a shank at a radial inner end of the airfoil body, theradially extending chamber extending at least partially into the shankto define a shank inner surface, wherein a first chordwise width of thehollow body at the tip end is less than a second chordwise width of thehollow body at the shank.